The present invention relates to aircraft gas turbine engines and, more particularly, to a fan blade containment structure for containing blade fragments ejected from damaged fan blades.
Gas turbine engines used in commercial aircraft are conventionally of a fan jet construction, i.e., the engines have a high volume fan at an axially forward end for forcing air into a first flow passage through an axial compressor, combustor and turbine and into a second bypass flow passage for providing thrust. The fan is positioned within a nacelle surrounding and spaced from a core engine housing containing the compressor, combustor and turbine, the space between the nacelle structure and core engine housing defining the bypass flow passage. During operation of the engine, and in particular during movement of an aircraft powered by the engine, the fan blades may be damaged by foreign objects such as, for example, birds or debris picked up on a runway. Impacts on the blades may damage the blades and result in blade fragments or entire blades being dislodged and flying radially outward at relatively high velocity.
Typically, fan jet engines have a fan casing circumscribing the engine within the nacelle and a blade containment structure circumscribing the fan casing. The casing may include a radial inner liner of noise absorption material, such as honeycomb panels, for reducing noise generated by the engine. One form of blade containment structure is described in U.S. Pat. No. 4,534,698 issued Aug. 13, 1985 and assigned to the assignee of the present invention, the disclosure of such patent being hereby incorporated by reference.
While the containment structure of the aforementioned U.S. patent has been effective in particular engines to provide the necessary containment of blade fragments, the recent development of larger engines with higher bypass ratios has revealed blade failure modes different than those of prior engines. In particular, in engines having relatively shorter axial fan casing dimensions, fan blade fragments have been found to be thrown radially outward and axially forward of the fan casing striking the inlet area at greater velocity than previously experienced. The resulting high energy impacts on the inlet inner liner may be sufficiently large to cause collapse of the acoustic honeycomb liner by compression of the honeycomb cell structure. Blade fragments may then exit tangentially through the inlet and, if the aircraft is in flight, perhaps result in damage to the aircraft. Accordingly, there is a need for a blade containment structure which can contain blade fragments ejected forward of a fan casing.